1. Field of the Invention
The invention relates to a control moment gyroscope and, more specifically, to a control moment gyroscope integral isolator which attenuates induced disturbance vibrations therein due to rotor static and dynamic unbalance and spin bearing anomalies.
2. Description of the Prior Art
The gyroscopic control apparatus generally referred to as a control moment gyroscope or CMG functions as the actuator or torquing element in an attitude control system of spacecraft and orbiting vehicles. The CMG includes an outer gimbal ring fixedly attached to the spacecraft and which includes a gimbal torque motor. The gimbal torque motor creates a torque about the gimbal axis which causes the CMG's gimbal to rotate about an axis perpendicular to the rotor spin axis. The rotation of the gimbal, moreover, creates a gyroscopic torque that is applied to the spacecraft structure for attitude control of the spacecraft.
In a vehicle having an ultra-precise pointing capability requirement, such as in a space telescope where precision pointing is mandatory and where vibration from any source will have a profound effect on the quality of the output image thereof, vibratory disturbances originating in the control moment gyroscope are of significant concern. That is, it is necessary to reduce the CMG's induced disturbance vibrations, because of their adverse effects upon spacecraft payload operation, to a very low level in order to satisfy some spacecraft and/or satellite attitude control system's pointing accuracy requirements. In the past, and as described in Applicant's assignees U.S. Pat. No. 3,918,778, significant progress has been made in the reduction of the principle vibration forcing functions, rotor and spin bearing unbalance forces. However, the present state of the art of rotor balance techniques and spin bearing design has not been able to reduce the disturbance vibrations to the levels required to meet the induced vibration requirements of an ultra-precise pointing spacecraft.
Accordingly, if an economical, generally vibrationless spacecraft, capable of ultra-precise pointing is to be developed, and as it is impractical to further reduce the vibratory forcing functions mentioned above, it is necessary to attenuate or isolate the disturbances transmitted from the gyroscopic rotor to the spacecraft. That is, some form of spring/damper must isolate the vibration generating elements from the spacecraft. Moreover, a control moment gyroscope attenuation or isolation system must be structurally sound and rigid along the spin axis and about the gimbal axis, and also satisfy the constraint that the torsional stiffness about the gyroscopic output axis remains high in order that an apparent increase in the CMG gimbal inertia, which would decrease the CMG's performance capability or increase the size of the CMG's gimbal torquer to maintain the same performance, is precluded.